NOVEL THERMAL METHOD FOR RAPID COKE MEASUREMENT IN LIQUID ROCKET ENGINES
There is disclosed a method of cleaning a liquid hydrocarbon-cooled bipropellant rocket engine. In an embodiment, method includes heating the engine to a temperature. The method includes applying ozone for a period of time. The method includes determining the temperature and the period of time are each sufficient to remove carbonaceous deposits. In another embodiment, the method may further include thermally imaging the heat transfer surfaces. There is disclosed apparatus for cleaning a liquid hydrocarbon-cooled bipropellant rocket engine. In one embodiment, the apparatus includes a heater, an ozone source, and a thermal camera. Other embodiments are also disclosed.
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This application claims the benefit under 35 U.S.C. 119 (e) of U.S. Provisional Patent Application No. 61/628,914, filed Nov. 10, 2011, by David Thomas Wickham, et al. for “A Novel Thermal Method for Rapid Coke Measurement in Liquid Rocket Engines,” which patent application is hereby incorporated herein by reference.
GOVERNMENT LICENSE RIGHTSThis invention was made with government support under contract number FA9300-12-M-1008 awarded by the United States Air Force. The government has certain rights in the invention.
BACKGROUNDGenerally, carbonaceous deposits (“coke”) mapping and removal from liquid hydrocarbon-fueled rocket engines is not undertaken as current launch systems dispose of the rocket engines after a single use. The NASA Space Transportation System, referred to as the Space Shuttle program, employed reusable rocket engines for main propulsion. However, these reusable rocket engines burned liquid hydrogen as fuel instead of a liquid hydrocarbon. Since it is impossible to form coke from liquid hydrogen, coke formation in the engine liner cooling passages was never an issue for the Space Shuttle program. No truly reusable liquid hydrocarbon-fueled rocket engine has yet been developed, and no reusable rocket launch system using liquid hydrocarbon fuels currently exists.
SUMMARYThis Summary is provided to introduce a selection of concepts in a simplified form that are further described below in the Detailed Description. This Summary is not intended to identify key aspects or essential aspects of the claimed subject matter. Moreover, this Summary is not intended for use as an aid in determining the scope of the claimed subject matter.
In an embodiment, there is provided a method of cleaning a liquid hydrocarbon-cooled bipropellant rocket engine, the method comprising heating the engine to a given temperature; applying ozone to flow through channels in the engine wall for a period of time; and determining the given temperature and the period of time are each sufficient to remove carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine.
In another embodiment, there is provided a method of cleaning a liquid hydrocarbon-cooled bipropellant rocket engine, the method comprising heating the engine to a given temperature; applying ozone to flow through channels in the engine wall for a period of time; and thermally imaging the heat transfer surfaces to determine the given temperature and the period of time are each sufficient to remove carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine.
In yet another embodiment, there is provided apparatus for cleaning a liquid hydrocarbon-cooled bipropellant rocket engine, the apparatus comprising a heater configured to heat the engine to a given temperature; an ozone source configured to apply a flow of ozone to channels in the engine wall for a period of time; and a thermal camera configured to thermally image the heat transfer surfaces to determine the given temperature and the period of time are each sufficient to remove carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine.
In still another embodiment, there is provided a method of cleaning a liquid hydrocarbon-cooled bipropellant rocket engine, the method comprising heating the engine to a given temperature; applying ozone to flow through channels in the engine wall for a period of time; and monitoring at least one of carbon dioxide and water emitted from the channels to determine the period of time are each sufficient to remove carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine.
In another embodiment, there is provided a method of cleaning a liquid hydrocarbon-cooled bipropellant rocket engine, the method comprising heating the engine to a given temperature; applying ozone to flow through channels in the engine wall for a period of time; and monitoring infrared thermal imaging of the engine to determine the period of time are each sufficient to remove carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine.
In yet another embodiment, there is provided apparatus for cleaning a liquid hydrocarbon-cooled bipropellant rocket engine, the apparatus comprising a heater configured to heat the engine to a given temperature; and an ozone source configured to apply a flow of ozone to channels in the engine wall for a period of time.
Additional objects, advantages and novel features of the technology will be set forth in part in the description which follows, and in part will become more apparent to those skilled in the art upon examination of the following, or may be learned from practice of the technology.
The patent or application file contains at least one drawing executed in color. Copies of this patent or patent application publication with color drawing(s) will be provided by the Office upon request and payment of the necessary fee.
Non-limiting and non-exhaustive embodiments of the present invention, including the preferred embodiment, are described with reference to the following figures, wherein like reference numerals refer to like parts throughout the various views unless otherwise specified. Illustrative embodiments of the invention are illustrated in the drawings, in which:
Embodiments are described more fully below in sufficient detail to enable those skilled in the art to practice the system and method. However, embodiments may be implemented in many different forms and should not be construed as being limited to the embodiments set forth herein. The following detailed description is, therefore, not to be taken in a limiting sense.
In an embodiment, there is provided a system and a procedure for the identification and simultaneous removal of carbonaceous deposits (“coke”) on heat transfer surfaces of liquid hydrocarbon-cooled bipropellant rocket engines. This technique involves a dilute mixture of ozone in oxygen and an inert gas, such as nitrogen. After warming the rocket thrust chamber and other fuel-cooled hardware to a temperature between 100 and 200° C., flowing this dilute ozone mixture through the engine liner cooling channels allows oxidation of the surface-adhered carbonaceous material to CO2 and water. In various embodiments, the ratio of ozone and oxygen to one or more inert gasses may include a wide range of percentages. Furthermore, ozone and oxygen may be applied to a rocket engine with or without other gasses, including inert gasses. This gentle, low temperature oxidation process releases heat that is conducted into the underlying metal (which is generally copper), allowing visualization of the areas where oxidation is taking place by monitoring the inner surface of the liner using a thermal imaging camera. Even though the diffusion of heat within the liner tends to smear out the heat release location, a high sensitivity thermal camera can still localize the area of heat release by observing temperature changes and differences on the order of 0.1° K. The removal of this thin layer of carbonaceous material (typically on the order of 1 to 2×10−6 inches thick) after each run allows the reuse of liquid hydrocarbon-fueled rocket engines by preventing the buildup of insulating coke layers from run to run. Coke has a much lower thermal conductivity than copper and thicknesses of only several millionths of an inch (e.g., about 0.0254 microns to about 0.0762 microns) can cause the wall temperatures to reach levels where they will fail causing loss of coolant which can also overheat the channel downstream of the leak, potentially causing skewed temperature profiles, low thrust performance, an early engine shutdown, or even complete engine failure. Knowing the location of coke deposits is helpful in that the quantity and location constitute information about the health of the engine and allows the operator to track changes in performance to estimate the remaining service life of the engine.
Carbon deposition in rocket engine cooling channels is potentially a very serious problem because current estimates suggest that coke has a thermal conductivity of only about 0.07 Btu/h-ft-° F. (0.12 W/m-C°) (approximately 3000 times lower than the copper of the thrust chamber walls.) (Edwards, J. T., “Liquid Fuels and Propellants for Aerospace Propulsion: 1903-2003”, AIAA Journal of Propulsion and Power, 19(6), 1089-1107 (2003)) The accumulation of even very thin layers of coke can cause the walls to reach temperatures where they will fail.
The location and thickness of coke layers deposited in copper test sections may be characterized by monitoring the outer surface temperature of a test article with a thermal imaging camera while the coke inside the test section is oxidized at low temperature to CO2 using a mixture of O3 in O2 and N2. The oxidation of coke with O3 produces 161 kcal of thermal energy for every gram mole of carbon oxidized:
C+2O3→CO2+2O2 ΔH=−162 kcal/gmol C Eq. 1
For example, it may be assumed that the coke layer is deposited on the three sides of the channel where the temperatures are hottest, i.e., the surface next to the hot gas of the thrust chamber and the two side walls of the channel, and that the channel width is 0.127 cm (0.050″) and the height is 0.381 cm (0.150″). The total surface area of the three sides in a one cm length is 0.889 cm2. If the coke layer is 2.54×10−6 cm (1.0×10−6 inches) thick, the total quantity of carbon present is 2.26×10−6 cm3. Using a coke density (Hazlett, R. N., “Oxidation Stability of Aviation Turbine Fuels,” ASTM, Philadelphia, Pa., (1991)) of 1.0 g/cm3, the deposit weight is 2.26×10−6 g, which is equivalent to 1.88×10−7 gram-moles of carbon. With the above heat of reaction, that quantity of coke would produce an exotherm of 0.031 calories or 0.128 joules when oxidized.
The magnitude of the temperature change produced in the length of the channel depends on the quantity of copper that bounds the three sides.
An illustration of a test apparatus 30 used to conduct experiments is shown in
Three processes were used to apply a layer of carbon to the internal surface of the copper test articles; The first process simply trapped a small sample of JP-7 fuel in the test section, which was placed in a muffle furnace at 400 to 450° C. for several hours. While simple, this method did not allow a uniform control or location of the coke deposits. The second process was used graphite paint as the composition was known an control of deposition occurred. Importantly, it is unknown as to the differences in ozone reactivity or graphite paint compared to coke. The third process included thermally cracking flowing n-dodecane or RP-2 fuel at high temperature and pressure. This method allowed deposition of coke in a manner more like the actual rocket engine coke deposits. The deposits could not be controlled for location of information or uniformity. Also, in early tests the copper tube was simply insulated on the back side and not heated independently. As a result, the temperature of the test section decreased from inlet to exit due to the transfer of heat to the ambient surroundings.
Phase I Test Results
Tubular Test Section
These factors explain the data of graph 50 shown in
Since the ozone generation process is not 100% efficient, turning on the O3 generator also results in some heating of the O2 flow. This shows up as an additional transient temperature effect during decoking. In order to eliminate this effect, the procedure may be changed to repeat the run after a decoking cycle to obtain a blank, allowing elimination of the effect of the O3 generator heating by subtraction. This process is shown in
Additional tests were also carried out with the graphite paint. Less overall carbon was applied and the paint was also added to both ends of the tubes in order to verify that the method has the potential to resolve the position of coke deposits down the length of the tube. The results of these tests are presented in
Overall these experiments demonstrate the fundamental feasibility of this method and the effectiveness of the data reduction routine. In addition, these results show the method may quickly identify locations that contain carbon deposited in the flow path, as well as showing that the carbon deposits have been removed.
Multi-Channel Plate (MCP) Test Section
Referring now to
During the low temperature oxidation runs, heat was applied with a flat electric heater that was attached to the back side of the MCP 80. The front side 86 of the multichannel plate 80, which has a thickness of 0.035″ (0.9 mm) to simulate the thrust chamber liner, faced the thermal imaging camera as shown in
In the first test we carried out on the MCP 80 to verify the heating and data reduction routine, an oxidation test was performed on the unit 80 without intentionally adding carbon to any of the channels in the plate. Therefore, the state of the test article was effectively “as received” from the machining and brazing processes. In this test, O2 was flowed at a rate of 3 standard liters per minute (slpm) through eight of the nine channels 82 in the plate 80 (although there are nine channels on the plate, space limitations prevented inclusion of all nine channels 82 in the flow manifolds).
The resulting temperature differences at three different times are plotted in
The CO2 produced was monitored during each test. During the initial oxidation, a measurable CO2 peak was obtained indicating that 0.098 mg of carbon had been oxidized, while during the subsequent baseline test we did not obtain any detectable level of CO2. Since any carbon was not intentionally added to the plate before this test, it was concluded that there was probably some carbon-containing material introduced into the plate during or after the face machining process and that the oxidation of this material probably caused the temperature increase at the front of the plate. On the other hand, since there was no measurable rise in the temperature at the back end of the plate, we conclude that no carbon-containing compound was located in this part of the plate.
After completion of this test, a series of tests were carried out to characterize the O3 lifetime in the plate after the carbon had been removed and found that a flow of about 2.5 slpm was required through each channel to prevent excessive O3 loss. Therefore, in subsequent tests, we increased the overall O2 flow to 10 slpm (the maximum flow for our mass flow controller) and only flowed O2 through a maximum of four channels during the oxidation step. While in a full scale system all channels would be oxidized together, the only effect this modification had was to reduce the sensitivity of our method.
For the next decoking test, carbon was deposited in two of the channels using graphite paint. To accomplish this the paint solution was diluted by a factor of 10 with acetone, resulting in a concentration of 1.5 mg graphite/ml solution. Then, approximately 0.2 ml of solution was injected into two of the channels in the plate, channels 4 and 6 (the two channels next to center channel) and flowed air through the channels until the acetone was completely evaporated. The total carbon weight added was 0.3 mg. In
In this figure, the temperature rise occurs primarily at the back end of the plate, indicating that the graphite was deposited in this end of the plate. In
The CO2 signal was integrated during the test and calculated so that it corresponded to the oxidation of 0.25 mg of carbon. This agrees well with the theoretical quantity of 0.3 mg initially deposited in the MCP. The thermal data show that the coke was located on the back half of the test section, therefore the average thickness of the graphite layer was approximately 2 microinches. Finally,
A second baseline run was carried out immediately after the first baseline in order to test the accuracy of our method. Since all of the carbon in the plate had been oxidized in the oxidation step, no carbon oxidation should be occurring in either the first or second baseline runs and therefore the temperature differences between the two runs should not show a change on one end as we saw previously in
After showing that detection of the oxidation of very thin layers of carbon in the multichannel plate, three tests were carried out in which we coked the inside of the multichannel test section with the new provisional RP-2 fuel. Heat was provided by heaters attached to both the front and back of the test section and the entire assembly was enclosed in mineral wool insulation. In all three tests, the test section was maintained at 28.2 atm (400 psig) and the front and back of the plate to 550° C. was heated as RP-2 was flowed through the channels at a rate of 1.0 ml/min. Overall, a cracking level of about 1% was obtained during this coking run.
The temperatures differences obtained by subtracting the baseline temperatures from those obtained during oxidation are shown in
Full Scale Decoking System Design
The Phase I results showed that our coke mapping and removal approach is very effective on laboratory scale hardware. To be of practical value, however, application to a full scale rocket engine may be made without excessive cost or effort. Perhaps the most challenging aspect is warming the hardware to temperatures between 150 to 200° C. where the low temperature oxidation process can occur. This step may be done rapidly to meet short turnaround times. An engineering analysis estimates the heating requirements and thermal response times of an engine. An assembly including a blower and a portable gas-fired heater, equivalent to the size of an outdoor gas grill, produces the necessary temperature in an engine in about 15 minutes and consumes only about five pounds of propane.
Rocket engines can be fairly massive objects, particularly those engines used as main propulsion on a space launch vehicle. Typical booster engine thrust levels are on the order of 300,000 to 1,000,000 pounds (136,000 to 454,000 kg.) Thrust-to-weight ratios of high performance engines are generally on the order of 100:1. An engine with a thrust of 750,000 lbf (340,194.3 kgf) would therefore weigh approximately 7500 lbm (3401.9 kgm). About ¼ of this weight is typically found in the thrust chamber and the regeneratively cooled section of the nozzle, which is where the fuel cooled passages that are prone to coking are located. Without knowing the specifics of an engine geometry, for the purposes of this heat transfer analysis the thrust chamber mass can be roughly assumed to be distributed as a simple annulus 1400, having a length L, a diameter Do, and a wall thickness t, as shown in
Using a full-scale engine cooled length of 70 inches (177.8 cm) and a thrust chamber wall thickness of 1 inch (2.54 cm) with an average diameter of 25 inches (63.5) yields an approximate mass of ˜1800 pounds (816.5 kg) of copper and nickel that would need to be heated for a decoking cycle. Using a specific heat of 0.093 Btu/Ibm-° F. (388 J/kg-° C.) for copper results in a total thrust chamber thermal capacitance mC of 175 Btu/° F. (332 kJ/° C.). This requires a minimum of 49,000 Btu's (51,650 kJ) to raise the temperature of the thrust chamber from 21 to 175° C. (where ozone oxidation of carbon works well) without any heat losses. Heating should probably be done within about 15 minutes to allow rapid ground processing of a reusable vehicle, therefore requiring a minimum heat transfer rate of 200,000 Btu/hr (58.6 kW) to the thrust chamber. This is about the same as two home central air furnaces. Pure electrical heating for a thermal load this size would be both expensive and taxing of many electrical supply installations, especially with multiple engines on a single vehicle. Even the ideal minimum heat load therefore suggests that heating an engine 1500 from inside along heat path 1502 using something like a ‘flameless’ radiant heater or blower with a propane burner, as shown in
Once the engine has been raised to the desired decoking temperature, it may be maintained for a period of time while temperatures stabilize, as well as during coke removal and thermal imaging. One good way to accomplish this is to use an electrically heated insulating blanket or mantle 1506 wrapped around the outside of the engine, as shown in
The engine may be modeled as a lumped thermal mass to estimate the size of the heater needed. Using convective heating and cooling to the inside and outside environments and a uniform internal temperature provided the governing linear, non-homogeneous ordinary differential equation (ODE):
Where T is the temperature of the rocket engine, mC is the engine thermal capacitance, U∞ is the effective external loss heat transfer coefficient, Ao is the external surface area, hi is the internal heat transfer coefficient, Ai is the engine internal combustion chamber and cooled nozzle surface area, qelec is the electrical heat input rate, Tgas is the temperature of the hot gas source used to heat the engine, and T∞ is the ambient temperature where the engine is being processed. This ODE may be used to immediately obtain some information about the heating process in the form of the time constant, τ:
Using an effective external loss heat transfer coefficient of 1.4 Btu/ft2-hr-° F., an external surface area of 44 ft2, and an internal heat transfer coefficient of ˜20 Btu/ft2-hr-° F. (i.e., turbulent air convection) with a surface area of ˜35 ft2 gives a time constant of about 14 minutes, which is actually surprisingly short. The values in the denominator indicate that this is mostly controlled by the convective heating process inside of the engine. The simple closed-form solution to the above ODE is:
Where T(t) is the thrust chamber temperature at time t, t/t is the non-dimensional time, and Teq is the equilibrium engine temperature attained if the heating process were continued over many time constants (>4). This equilibrium temperature is given by:
The closed form solution can be easily inverted to determine the average internal engine gas temperature required to meet the 15 minute heating goal to 175° C., yielding Teq=253° C. and Tgas=247° C. (assuming a 10 kW total electrical heat input rate). This average gas temperature would seem to be reasonable, and not likely to overheat the hardware, and it should also allow volatilization and driving out any trapped fuel from the engine cooling passages during heating. This solution corresponds to an initial hardware heat input rate of 304,000 Btu/hr (89.0 kW) that could be satisfied with a low pressure rise air blower of about 500 cfm (14150 slpm) capacity (i.e., ˜¼ horsepower (0.19 kW)), and a total propane flow of about 5 pounds (2.3 kg) to accomplish the heating of a single engine.
This analysis therefore indicates that it would be feasible to heat an engine in 15 minutes, allow its temperature to stabilize over the next 15 minutes, perform the ozone decoking and coke mapping over the next 15 minutes under constant input power, then remove the external heating blanket and allow the engine to cool while continuing to circulate cool air inside with the blower, as indicated in
The total quantity of carbon present can be calculated by monitoring the CO2 concentration in the effluent during ozone oxidation. Even with the thin carbon layers deposited, online analyzers like the California Analytical NDIR CO2 have adequate sensitivity to provide the continual measure of total coke oxidation rate. A coke thickness of 1.0×10−6 inches was calculated above with 2.26×10−6 grams or 1.88×10−7 moles carbon in a 1 cm (0.39-in) length of a single channel. If a similar thickness is assumed through the entire cooling channel flow path, 60 inches long and 400 cooling channels, the total quantity of coke is estimated to be 0.138 grams of carbon or 0.0114 gram moles. When oxidized, 0.0114 moles of carbon will produce a total volume of 0.276 standard liters of CO2. The total flow area in all channels is 19.2 cm2 (0.048 cm2 per channel×400 channels) and therefore if the total gaseous flow rate is 5 slpm, the velocity through the channels is 258 cm/min (102 in/min) resulting in a residence time of less than a minute, which will provide an accurate picture of the amount of carbon present in the cooling channels. Finally, if all of the carbon is assumed to be oxidized to CO2 in a period of 15 min, the average CO2 concentration will be about 0.37%. Online analyzers are available with detection limits of 0.0001% or 1 ppm, so concentrations of 0.37% (3700 ppmv) can be readily detected.
Kinetics of Coke Oxidation
Previous work (Wickham, D. T., Engel, J. R., Jones, M. and Windecker, B., “Methods to Remove Coke from Endothermic Heat Exchangers,” NASA Phase II Final Report, Contract No. NNCO5CAO5C (2007)) showed that the rate of coke oxidation at temperatures between 100 and 200° C. was very rapid, even with layers approaching thicknesses of thousandths of an inch. In most cases the oxidation was complete in periods of several minutes. Thus, the CO2 concentrations in the effluent will be higher than those estimated above. In addition, it is possible that the rate of coke oxidation will be determined by the flux of O3 to the carbon surface. At a velocity of 258 cm/min, and with an O3 concentration of 0.5%, mass transfer calculations show that all the O3 contained in the flow will contact the surface layers of carbon within a distance of several centimeters and, therefore, it is possible that all of the O3 would be consumed in a relatively short distance. In this case, a temperature increase would be measured only in the first part of the flow path. However, as the carbon in that section is consumed, more O3 is available for the next portion of the reactor so that an oxidation front moves down the length of the cooling channels. With scan rates of up to 60 Hz, the FLIR SC325 can easily capture this movement of the oxidation front. In this case, the concentration of CO2 exiting the manifold is determined by the reaction stoichiometry shown in Eq. 1 and the O3 concentration entering the manifold. Moreover, the thickness of the layer at any position can be determined by the magnitude and duration of localized temperature increases. Once the IR camera shows that the oxidation front has reached the end of the flow path, the production of CO2 stops.
Hazard Mitigation
Our coke mapping and removal process can be carried out very safely because the two potential hazards: 1) introducing an oxidizing agent into the fuel cooling channels, and 2) exposure to O3, can be effectively mitigated. To prevent hazards associated with flowing O3 or O2 through the cooling channels, the first step in the process is to flush the channels with N2 at ambient temperature removing any liquid fuel left in the heat exchanger. A purge is continued as a heater is then used to raise the heat exchanger temperature to about 175° C., which will vaporize the last residuals of fuel. Moreover, when low concentrations of O2 and O3 are introduced, the system will be operating at atmospheric pressure and the quantities of coke are small enough that very little temperature rise can occur from the oxidation process. Finally, after the process has been completed the fuel channels can again flushed with N2 to remove any traces of O2 before the vehicle is refueled.
Hazards associated with O3 exposure are prevented by several factors. First, the 0.1 to 0.5% O3 levels used are generated as needed by directing an O2 flow through an online O3 generator, eliminating any requirement to store or handle O3. Next, while O3 is flowing through the channels, all flows exiting the manifold are directed through a commercial O3 decomposition catalyst before it is vented to atmosphere. The catalyst modules are 99.9% efficient and cost less than $50/unit. (Absolute Ozone, accessed at http://www.ozonesolutions.com/products/Ozone-Destruct-Units/ODS-3_High_Flow_Ozone_Destruct_Unit, 2011) If the concentration exiting the heat exchanger is 0.15% using two catalyst beds in series would produce a concentration 0.0015 ppm, a factor of 66 less than the 0.1 ppm OSHA 8 hour PEL. (Occupational Safety and Health Administration, “Limits for Air Contaminants,” Code of Federal Regulations 29 CFR 1910, Subpart Z, Table Z-1. Accessed at: http://www.osha.gov/pls/oshaweb/owadisp.show_document?p_table=STANDARD S&p_id=9992) Moreover, once the flow is diluted in the atmosphere the concentrations will be even lower. O3 has been used without incident in a number of projects, including the development of methods to oxidize coke deposits in endothermic heat exchangers and to control waste streams produced in space missions.
Advantages of this Technology
The low temperature oxidation approach has a number of advantages over other potential methods to map coke deposition. First of all, it is a rapid, non-intrusive way to characterize coke deposition over the entire cooling channel assembly. It provides accurate, reliable data to be obtained that then allows engine life prediction models to be generated and to determine required service intervals. It can be done quickly while the engine is still on the vehicle. Finally, it is a safe and simple method that does not require handling or storage of any hazardous chemicals. The gentle, low temperature oxidation process allowed by the use of ozone results in nearly zero oxidation of the underlying metal. While coke removal can also be done with oxygen, the high temperatures required (>500° C.) can cause significant oxidative attack on the metal as well as warpage and destruction of elastomers.
In addition to all of the attractive features outlined above, this method has another significant advantage. In the process of mapping the coke deposition, it completely removes all of the coke from the heat exchanger pathways. Therefore, it solves the difficult and challenging problem of coke removal from a complex flow path. Moreover, the mapping process does not have to be employed every time coke is removed from the cooling channels. One scenario would be to carry out the imaging with coke oxidation at selected times, for example during engine development or when the engine has completed a specific number of missions. However, the low temperature ozone oxidation process could be used without the camera as a part of the routine maintenance between each mission, to remove any coke that accumulated during the previous mission. This would save time and reduce cost, but still allow the engine to be reused safely. In addition, the total quantity of coke removed in each treatment could be determined by attaching a CO2 analyzer to the lines exiting the channels. Thus, there are several ways this technology could be employed to support the particular engine maintenance or development goals.
Alternative Ways to Make or Use this Technology
It is possible to operate the ozone coke removal process without employing the thermal camera for mapping of the deposit location. It may also be possible to identify another active oxidation compound that could fulfill the role of ozone in this invention. More active oxidizers could potentially lower the hardware preheating requirement, which would also improve the baseline hardware temperature stability and therefore improve the sensitivity of this method. Removing coke with oxygen is also possible, but requires temperatures on the order of 500° C. (932° F.) to accomplish. Many components cannot withstand temperatures this high due to the presence of elastomers, and significant oxidation of surfaces can occur along with dimensional changes or warpage.
Limitations
It is possible for the coke mapping process to not work if the coke layers are too thin or if the thermal mass and thermal conductivity of the thrust chamber is too large for the heat from coke oxidation to increase the temperature of the liner more that the Noise Equivalent Temperature Difference (NETD) of the thermal camera. Another limitation on detection level occurs if insufficient time is given for the rocket engine to approach a steady-state temperature distribution. While an engine does not need to be completely isothermal, engine heating transients must be allowed to decay before coke mapping can be accomplished. In addition, this steady-state temperature distribution must be maintained to the end of the second oxidation cycle in order to identify the small temperature differences associated with coke oxidation.
Although the above embodiments have been described in language that is specific to certain structures, elements, compositions, and methodological steps, it is to be understood that the technology defined in the appended claims is not necessarily limited to the specific structures, elements, compositions and/or steps described. Rather, the specific aspects and steps are described as forms of implementing the claimed technology. Since many embodiments of the technology can be practiced without departing from the spirit and scope of the invention, the invention resides in the claims hereinafter appended.
Claims
1. A method of cleaning a liquid hydrocarbon-cooled bipropellant rocket engine, the method comprising:
- heating the engine to a given temperature;
- applying ozone to flow through channels in the engine wall for a period of time; and
- determining the given temperature and the period of time are each sufficient to remove carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine.
2. The method of claim 1 wherein the step of heating the engine to the given temperature comprises the given temperature in a range from 150° C. to 200° C.
3. The method of claim 1 wherein the step of heating the engine to the given temperature comprises the given temperature in a range from 160° C. to 180° C.
4. The method of claim 1 wherein the step of applying ozone to flow through channels in the engine wall for a period of time comprises the period of time within a range of 10 to 60 minutes.
5. The method of claim 1 wherein the step of applying ozone to flow through channels in the engine wall for a period of time comprises the period of time within a range of 10 to 20 minutes.
6. The method of claim 1 wherein the step of applying ozone to flow through channels in the engine wall for a period of time comprises the period of time within a range of about 15 minutes.
7. The method of claim 1 wherein the step of determining the given temperature and the period of time are each sufficient to remove the carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine comprises maintaining the engine at the given temperature and flowing the ozone through the channels for the period of time based on a size of the engine.
8. The method of claim 1 wherein the step of determining the given temperature and the period of time are each sufficient to remove the carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine comprises maintaining the engine at the given temperature and flowing the ozone through the channels for the period of time based on a size of the engine.
9. The method of claim 1 wherein the step of determining the given temperature and the period of time are each sufficient to remove the carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine comprises maintaining the engine at the given temperature and flowing the ozone through the channels for the period of time based on thermal imaging of the heat transfer surfaces.
10. The method of claim 1 further comprising placing an insulating jacket on an outside surface of the rocket engine so as to maintain the engine at the given temperature.
11. The method of claim 1 further comprising placing an electric heating mantle on an outside surface of the rocket engine so as to maintain the engine at the given temperature.
12. The method of claim 1 further comprising placing an insulating jacket and an electric heating mantle on an outside surface of the rocket engine so as to maintain the engine at the given temperature.
13. The method of claim 11 further comprising placing a heating manifold insert within an interior region of the rocket engine so as to maintain the engine at the given temperature.
14. The method of claim 1 further comprising placing a heating manifold insert within an interior of the rocket engine so as to maintain the engine at the given temperature.
15. The method of claim 1 further comprising directing a path of heat flow from a propane burner within an interior of the rocket engine so as to maintain the engine at the given temperature.
16. The method of claim 1 further comprising purging the engine with nitrogen to remove fuel prior to applying ozone.
17. The method of claim 1 further comprising thermally imaging the heat transfer surfaces.
18. The method of claim 17 wherein the thermal imaging of the heat transfer surfaces includes directing a thermal camera toward an interior region of the rocket engine so as to image the heat transfer surfaces of the rocket engine.
19. The method of claim 18 wherein the thermal imaging of the heat transfer surfaces includes a mirror guide disposed within the interior region of the rocket engine, a mirror selectively moveable along the mirror guide, and an actuator to reposition the mirror along the mirror guide at different distances from the thermal camera with the thermal camera being directed toward an interior region of the rocket engine so as to image the heat transfer surfaces of the rocket engine.
20. The method of claim 19 further comprising processing a position of the mirror with respect to the mirror guide so as to determine a position of an imaged location of the heat transfer surfaces of the rocket engine.
21. A method of cleaning a liquid hydrocarbon-cooled bipropellant rocket engine, the method comprising:
- heating the engine to a given temperature;
- applying ozone to flow through channels in the engine wall for a period of time; and
- thermally imaging the heat transfer surfaces to determine the given temperature and the period of time are each sufficient to remove carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine.
22. The method of claim 21 wherein the thermal imaging of the heat transfer surfaces includes directing a thermal camera toward an interior region of the rocket engine so as to image the heat transfer surfaces of the rocket engine.
23. The method of claim 22 wherein the thermal imaging of the heat transfer surfaces includes a mirror guide disposed within the interior region of the rocket engine, a mirror selectively moveable along the mirror guide, and an actuator to reposition the mirror along the mirror guide at different distances from the thermal camera with the thermal camera being directed toward an interior region of the rocket engine so as to image the heat transfer surfaces of the rocket engine.
24. The method of claim 23 further comprising processing a position of the mirror with respect to the mirror guide so as to determine a position of the thermal camera with respect to an imaged location on the heat transfer surfaces of the rocket engine.
25. The method of claim 23 further comprising purging the engine with nitrogen to remove fuel prior to applying ozone.
26. A method of cleaning a liquid hydrocarbon-cooled bipropellant rocket engine, the method comprising:
- heating the engine to a given temperature;
- applying ozone to flow through channels in the engine wall for a period of time; and
- monitoring at least one of carbon dioxide and water emitted from the channels to determine the period of time are each sufficient to remove carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine.
27. The method of claim 26 further comprising monitoring infrared thermal imaging of the engine to determine the period of time are each sufficient to remove carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine.
28. The method of claim 26 further comprising purging the engine with nitrogen to remove fuel prior to applying ozone.
29. A method of cleaning a liquid hydrocarbon-cooled bipropellant rocket engine, the method comprising:
- heating the engine to a given temperature;
- applying ozone to flow through channels in the engine wall for a period of time; and
- monitoring infrared thermal imaging of the engine to determine the period of time is sufficient to remove carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine.
30. The method of claim 29 further comprising monitoring carbon dioxide emitted from the channels to determine the period of time is sufficient to remove carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine.
31. The method of claim 29 further comprising mapping coke deposits based on the step of monitoring infrared thermal imaging to determine where the coke deposits are located in the channels in the engine wall.
32. The method of claim 29 further comprising purging the engine with nitrogen to remove fuel prior to applying ozone.
33. The method of claim 29 further comprising purging the engine with nitrogen to remove fuel prior to applying ozone.
34. Apparatus for cleaning a liquid hydrocarbon-cooled bipropellant rocket engine, the apparatus comprising:
- a heater configured to heat the engine to a given temperature; and
- an ozone source configured to apply a flow of ozone to channels in the engine wall for a period of time.
35. Apparatus according to claim 34 further comprising a thermal camera configured to thermally image the heat transfer surfaces to determine the given temperature and the period of time are each sufficient to remove carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine.
36. Apparatus according to claim 34 further comprising a carbon dioxide detector configured to monitor carbon dioxide emitted from the channels to determine the period of time are each sufficient to remove carbonaceous deposits from heat transfer surfaces of the liquid hydrocarbon-cooled bipropellant rocket engine.
37. Apparatus according to claim 34 further comprising a nitrogen source configured to purge the engine with nitrogen to remove fuel prior to applying ozone.
38. The apparatus of claim 34 wherein the thermal camera has a lens directed toward an interior region of the rocket engine so as to image the heat transfer surfaces of the rocket engine.
39. The apparatus of claim 38 further including a mirror guide disposed within the interior region of the rocket engine, a mirror selectively moveable along the mirror guide, and an actuator to reposition the mirror along the mirror guide at different distances from the thermal camera with the lens of the thermal camera being directed toward an interior region of the rocket engine so as to image the heat transfer surfaces of the rocket engine.
40. The apparatus of claim 39 further comprising a processor configured to receive a position of the mirror with respect to the mirror guide so as to determine a position of an imaged location of the heat transfer surfaces of the rocket engine.
41. The apparatus of claim 34 further comprising an insulating jacket configured to cover an outside surface of the rocket engine so as to maintain the engine at the given temperature.
42. The apparatus of claim 34 further comprising an electric heating mantle configured to cover an outside surface of the rocket engine so as to maintain the engine at the given temperature.
43. The apparatus of claim 34 further comprising an insulating jacket and an electric heating mantle configured to cover an outside surface of the rocket engine so as to maintain the engine at the given temperature.
44. The apparatus of claim 43 further comprising a heating manifold insert configured to position within an interior region of the rocket engine so as to maintain the engine at the given temperature.
45. The apparatus of claim 34 further comprising a heating manifold insert configured to position within an interior of the rocket engine so as to maintain the engine at the given temperature.
46. The apparatus of claim 34 further comprising a propane burner configured to direct a path of heat flow from within an interior of the rocket engine so as to maintain the engine at the given temperature.
47. The apparatus of claim 34 further comprising a source of nitrogen configured to purge the engine with the nitrogen to remove fuel prior to applying ozone.
Type: Application
Filed: Nov 13, 2012
Publication Date: Aug 8, 2013
Applicant: REACTION SYSTEMS, LLC (Golden, CO)
Inventor: REACTION SYSTEMS, LLC (Golden, CO)
Application Number: 13/675,786
International Classification: B08B 7/00 (20060101);